The present invention generally relates to generated air pressure differentials that can be used to actuate mechanical devices. More specifically, the present invention relates to centrifugally generated air pressure differentials by rotor blades that can be used to actuate active rotor blade control surfaces.
The tips of propeller blades of a propeller driven fixed wing aircraft rotate in a directional plane that is usually perpendicular to the direction of flight. In contrast, ends of main rotor blades of a helicopter usually rotate in a directional plane that parallels the direction of flight. The aerodynamic environment of rotating helicopter rotor blades is non-symmetric, unsteady and three-dimensional. Compressibility effects are experienced on an advancing rotor blade. Reverse flow and dynamic stall occur on a retreating rotor blade and a returning wake is encountered under certain flight conditions. The environmental effects on the rotating blades results in large blade vibrations that are filtered through to the fuselage. The fuselage vibration arising primarily from the main rotor blades limits helicopter performance and reduces the structural life of helicopter components, which presents maintenance expense and reliability issues. The high vibration encounter also leads to pilot fatigue and poor ride quality. Significant research is presently dedicated toward realizing the goal of smooth and quiet helicopter flight to improve the cost effectiveness and wider community acceptance of helicopters.
Passive vibration absorbers and isolators are one solution, but lead to a large weight penalty and exhibit poor off-design performance. Active and smart vibration reduction techniques using adaptive control strategies that employ sensors and actuators enable performance and ride quality improvement. Improved vibration reduction can be achieved by placing the actuators as close as possible to the vibration source, in this case the main rotor blades. Higher harmonic control of a fixed swashplate superimposes a multi-cyclic pitch input on the rotor blades, which is beyond what is imposed by the primary flight controls. The next step beyond fixed swashplate actuation is on-blade actuation using active control surfaces to directly alter the rotor blade aerodynamic loads and load distributions. On-blade active systems would open a new technique for vibration control, aeromechanical stability augmentation, handling qualities enhancement and noise reduction.
There are two rotor blade control surface concepts that are particularly attractive for active control. The first concept incorporates trailing edge flaps to affect vibration control, where piezoelectric devices are often proposed as the actuation mechanism. In this approach, each trailing edge flap is a discrete element of the rotor blade and is not part of the direct load path. One of the advantages of a discrete actuation system is that multiple flaps can be mounted on a blade, offering greater control flexibility. The flap actuator designs that incorporate piezoelectric bender actuators have a low force output and in general, are restricted to small-scale wind tunnel models. Bender designs utilizing layered structures, bimorphs, and wing spar cantilevers have been investigated for small flaps in the range of 1-5 inches. These small flaps have achieved ±2° to ±5° flap deflection and reduced vibration. However, the larger deflections were for smaller flaps and most of the results were under either low load or low speed conditions. Piezoelectric stack actuators, typically used in larger scale applications, have a larger force output than benders but produce a smaller displacement. Designs that have incorporated lever-fulcrum and “X-frame” type amplification schemes have been investigated resulted in ±6° to ±4° flap deflection for a 4 inch and a 7 inch flap, respectively. For a larger full scale flap of about 36 inches, a “double X”, X-frame type amplifier was also investigated. Flap deflections in the range of ±2-3° are expected. As of late, an induced shear PZT tube actuator was incorporated into a hinge amplification device to deflect full-scale rotor blade trailing edge flaps. Experimental bench top testing of a full-scale tube actuator flap system was conducted to validate the analysis. The experimental testing revealed that for an applied electric field of 3 kV/cm, the tube actuator deflected a representative full scale 12 inch flap ±2.8° at 0 RPM and ±1.4° for a 400 RPM condition.
The second concept for on-blade actuation that is being researched is the use of Miniature Trailing-Edge Effectors (MiTEs) on the rotor blades. MiTEs are an extension of the passive high-lift device, the Gurney flap. Gurney flaps are small flat plates, between 0.5 to 5 percent chord, fitted perpendicularly to the airfoil surface at or near the trailing edge of a wing or rotor blade. A MiTE is an active Gurney flap, which can be used to actively control the lift and moment distribution on a rotor blade. MiTEs also have the advantage of having very low actuator loads compared to those of traditional trailing-edge flaps. Recently, experimental and validated computational fluid dynamics research has been done on MiTEs and an unsteady aerodynamic model was created for MiTEs placed at the trailing edge. The model was modified to account for a MiTE placed at the trailing edge and up to the 85 percent chord position. The aerodynamic model was also incorporated into a rotor performance code software to predict the effect of MiTEs on rotor performance and explore the ability to extend the flight envelope of the RAH-66 Comanche. The maximum velocity of Comanche was shown to have the potential increase of 20 percent with the increased use of transonic airfoils as facilitated through the use of MiTEs on the outboard section of the rotor blades. Investigations have also been made on increasing the service ceiling of the Comanche, which showed a potential improvement of 8 percent with the use of MiTEs. MiTEs could be used in place of active flaps for vibration control, as the MiTEs appear to be ideal in providing the required changes in lift and moments required for individual blade control. An additional advantage of using MiTEs is that they provide this potential with significantly lower actuator loads and are insensitive to compressibility effects.
There exist several reasons why piezoelectric actuated on-blade surfaces are problematic. Piezoelectric material is quite heavy with a density of nearly 7500 kg/m^3, which is nearly equal to that of steel. In helicopters in particular, weight is a critical design parameter. In general, several pounds of extra weight must be added to a helicopter blade to incorporate a trailing edge flap. Because each individual actuator will add a significant amount of weight, providing redundancy in the flap actuation design will necessarily imply that the actuator weight will be at least doubled. In addition to the weight issue, piezoelectric material is also very brittle. Piezoelectric actuators often fail in service due to multi-axial loads on the rotor blades that are unintentionally applied because of their brittle nature. The high failure rate of piezoelectric actuators would most likely imply that actuator redundancy would be necessary, which again would result in a high weight penalty. In addition, the output force of piezoelectric materials is directly proportional to the applied voltage. Therefore, a high voltage (˜1000V) is required to obtain the full stroke of the actuator. Reliably supplying this high voltage out to the rotating frame of a helicopter is difficult to achieve through existing electric slip rings. Finally, piezoelectric stack actuators require significant structural housings to provide precompression and mechanical linkages to amplify extremely small piezoelectric strain of about 0.15%, which necessarily results in large weight penalties for on-blade actuation. A lighter, more reliable and more redundant actuation method would offer helicopter designers a more feasible and attractive option for on-blade actuation, which the current invention supplies.
Another method of actuating active blade surfaces could be through the use of electromechanical motors. Typically, electromechanical motors are geared to rotary outputs that have fixed amplitudes. Therefore, the ability to achieve varying dynamic flap amplitudes at different frequencies would be difficult, if not impossible. An alternative design might employ a jack screw geared to the motor output. The axis of the jack screw would be attached to the flap axis and the motor would need to reverse directions at high frequencies of about 20-30 Hz. This design would suffer from a possible undesirable failure mode, where the motor might fail when the flap is fully deflected, causing an unstable aerodynamic effect. Finally, employing an electromechanical motor with enough torque or force to deflect a full-scale trailing edge flap would result in a significant weight penalty.
It is an object of the present invention to provide actuation of devices on a rotor blade with air pressure generated by the rotor blade.
It is an object of the present invention to provide actuation of devices on a rotor blade with minimal connections off of the rotor blade.